Aircraft with a fuselage accommodating an unducted turbine engine

ABSTRACT

An aircraft comprising a fuselage and an unducted turbine engine. The fuselage having a divot with an upstream edge and a downstream edge. The divot is defined by a straight reference line having a length (L) and a maximum depth (h) relative to the straight reference line. The unducted turbine engine having an engine core, a nacelle, and a set of blades. A first flow ratio (FR1) is equal to: h/L.

TECHNICAL FIELD

The disclosure generally relates to an aircraft, and more specifically,to an aircraft having a fuselage that accommodates an unducted turbineengine.

BACKGROUND

Turbine engines are rotary engines that extract energy from a flow ofworking air passing serially through a compressor section, where theworking air is compressed, a combustor section, where fuel is added tothe working air and ignited, and a turbine section, where the combustedworking air is expanded and work taken from the working air to drive thecompressor section along with other systems, and provide thrust in anaircraft implementation. The compressor and turbine stages compriseaxially arranged pairs of rotating blades and stationary vanes. Theunducted turbine engine is arranged as an engine core comprising atleast a compressor section, a combustor section, and a turbine sectionin axial flow arrangement and defining at least one rotating element orrotor and at least one stationary component or stator.

Turbine engines come in different configurations, such as a turbopropengine, which is a turbine engine that drives an aircraft propeller, aturbofan engine, which is a turbine engine with a fan upstream of theengine core, with both the fan and the engine core being received withina nacelle, and a propfan turbine engine, which is also called anunducted turbine engine. The unducted turbine engine includes aspects ofboth turboprop engine and the turbofan engine. For example, an unductedturbine engine include a set of rotating blades, or propellers, on theexterior of the engine casing similar to a turboprop, without therotating blade being constrained within the nacelle. The lack of anacelle or other casing surrounding the rotating blades of the fansection, lead to the name of an “unducted” fan or propfan engine.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present description, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which refers to the appended FIGS., inwhich:

FIG. 1 is a schematic cross-sectional view of an unducted turbineengine.

FIG. 2 is a schematic perspective view of an aircraft including anunducted turbine engine suitable for use as the unducted turbine engineof FIG. 1 , with the unducted turbine engine mounted to an aircraftfuselage having a divot.

FIG. 3 is a top-down view of the aircraft as seen from a horizontalplane above the aircraft of FIG. 2 .

FIG. 4 is the same view of the aircraft as shown in FIG. 3 , and furtherillustrating a freestream airflow along with an inlet airflow duringoperation of the aircraft.

FIG. 5 is a schematic front view of an exemplary aircraft suitable foruse as the aircraft of FIG. 2 , viewed from a vertical plane in front ofthe aircraft, and further illustrating additional geometriccharacteristics of the aircraft.

FIG. 6 is a front view of an alternative aircraft suitable for use asthe aircraft of FIG. 2 , viewed from a vertical plane, furtherillustrating a V-Tail and additional geometric characteristics of theaircraft.

FIG. 7 is a side view of the aircraft of FIG. 6 , viewed from a verticalplane, further illustrating additional geometric characteristics of theaircraft.

DETAILED DESCRIPTION OF THE EMBODIMENTS

The present disclosure relates to an aircraft including a fuselage thatis shaped to accommodate an unducted turbine engine mounted directly tothe fuselage, such as the empennage. An unducted turbine engine hasexposed fan blades, which are not enclosed within an outer nacelle orfan duct

Historically, unducted turbine engines, while more fuel efficient atcommercial aircraft cruise speeds than ducted turbofan engines, have hadother undesirable characteristics, such as comparatively loud noiselevels

The space between the unducted turbine engine and the fuselage acts as anozzle that increases the local airflow speed through the exteriorlylocated fan blades, between the nacelle and the blade tips, withrelation to a freestream airflow upstream of the unducted turbineengine. The annulus of air entering the blades between the outer surfaceof the nacelle and a swept circle defined by the blade tips is thoughtof as the “inlet airflow” to the blades. This acceleration of the inletairflow to the blades negatively affects the fuel consumption of theunducted turbine engine. To counter, and even negate this acceleration,a divot is formed in the fuselage, at a location generally correspondingto the unducted turbine engine. The divot is shaped to control the speedof the inlet airflow such that the inlet airflow is the same or lessthan the freestream air speed. It has been found that by controlling theshape (depth, length, curvature) and location of the divot, independence on the type of unducted engine, one can advantageouslycontrol the local airspeed to improve engine performance.

Reference will now be made in detail to present embodiments of thedisclosure, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the disclosure.

The word “exemplary” is used herein to mean “serving as an example,instance, or illustration.” Any implementation described herein as“exemplary” is not necessarily to be construed as preferred oradvantageous over other implementations. Additionally, unlessspecifically identified otherwise, all embodiments described hereinshould be considered exemplary.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within anunducted turbine engine or vehicle, and refer to the normal operationalattitude of the unducted turbine engine or vehicle. For example, withregard to a unducted turbine engine, forward refers to a position closerto an engine inlet and aft refers to a position closer to an enginenozzle or exhaust.

As used herein, the term “upstream” refers to a direction that isopposite the fluid flow direction, and the term “downstream” refers to adirection that is in the same direction as the fluid flow. The term“fore” or “forward” means in front of something and “aft” or “rearward”means behind something. For example, when used in terms of fluid flow,fore/forward means upstream and aft/rearward means downstream.

The term “fluid” may be a gas or a liquid, or multi-phase. The term“fluid communication” means that a fluid is capable of making theconnection between the areas specified.

Additionally, as used herein, the terms “radial” or “radially” refer toa direction away from a common center. For example, in the overallcontext of a turbine engine, radial refers to a direction along a rayextending between a center longitudinal axis of the engine and an outerengine circumference.

All directional references (e.g., radial, axial, proximal, distal,upper, lower, upward, downward, left, right, lateral, front, back, top,bottom, above, below, vertical, horizontal, clockwise, counterclockwise,upstream, downstream, forward, aft, etc.) are only used foridentification purposes to aid the reader's understanding of the presentdisclosure, and do not create limitations, particularly as to theposition, orientation, or use of aspects of the disclosure describedherein. Connection references (e.g., attached, coupled, connected, andjoined) are to be construed broadly and can include intermediatestructural elements between a collection of elements and relativemovement between elements unless otherwise indicated. As such,connection references do not necessarily infer that two elements aredirectly connected and in fixed relation to one another. The exemplarydrawings are for purposes of illustration only and the dimensions,positions, order and relative sizes reflected in the drawings attachedhereto can vary.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise. Furthermore, as used herein, theterm “set” or a “set” of elements can be any number of elements,including only one.

“High” and “Low” as used herein are descriptors with regards to theperformance indicator quantities described herein.

As used herein, the term “Mach number” will refer to the speed of afluid around a body without the existing speed of the fluid. Forexample, in aviation an observed Mach number can be decreased orincreased based on the wind speed and direction of the wind that theaircraft is flying in. In cases where the wind is a tail wind (e.g.,flowing from tail to nose), the speed of the wind can be added tooverall speed or Mach number of the aircraft. In cases where there is ahead wind (e.g., flowing from nose to tail), the speed of wind can besubtracted to the overall speed or Mach number of the aircraft. The Machnumber, as discussed herein, is assuming no tail wind or head wind.

As used herein, the term “cruise speed” refers to operation of a turbineengine utilized to power an aircraft may operate at a cruising speedwhen the aircraft levels after climbing to a specified altitude. Aturbine engine may operate at a cruising speed that is from 50% to 90%of a rated speed, such as from 70% to 80% of the rated speed. In someembodiments, a cruising speed may be achieved at about 80% of fullthrottle, such as from about 50% to about 90% of full throttle, such asfrom about 70% to about 80% full throttle. As used herein, the term“cruise flight” refers to a phase of flight in which an aircraft levelsin altitude after a climb phase and prior to descending to an approachphase. In various examples, cruise flight may take place at a cruisealtitude up to approximately 65,000 ft. In certain examples, cruisealtitude is between approximately 28,000 ft. and approximately 45,000ft. In yet other examples, cruise altitude is expressed in flight levels(FL) based on a standard air pressure at sea level, in which cruiseflight is between FL280 and FL650. In another example, cruise flight isbetween FL280 and FL450. In still certain examples, cruise altitude isdefined based at least on a barometric pressure, in which cruisealtitude is between approximately 4.85 psia and approximately 0.82 psiabased on a sea-level pressure of approximately 14.70 psia and sea-leveltemperature at approximately 59 degrees Fahrenheit. In another example,cruise altitude is between approximately 4.85 psia and approximately2.14 psia. It should be appreciated that, in certain examples, theranges of cruise altitude defined by pressure may be adjusted based on adifferent reference sea-level pressure and/or sea-level temperature.

In certain exemplary embodiments of the present disclosure, an unductedturbine engine having a set of circumferentially spaced fan blades,which extend, exteriorly, beyond a nacelle encasing an engine core, isconnected to an aircraft fuselage by a pylon, which spaces the unductedturbine engine from the fuselage, and an associated tail wing, resultingin a channel between the fuselage/tail wing and the unducted turbineengine.

During operation of the aircraft (e.g., including a fuselage without thedivot as described herein) the freestream air flows over a portion ofthe fuselage and through the channel between the fuselage and theunducted turbine engine. At least a portion of the freestream airflowforms the inlet airflow, which flows through the annulus between theouter surface of the nacelle and the swept circle defined by the bladetips. The inlet airflow is ultimately used to generate a portion of anoverall thrust of the unducted turbine engine. The channel acts as anozzle that accelerates the inlet airflow.

The freestream airflow and the inlet airflow are both defined byrespective Mach numbers. In conventional aircraft, the acceleration ofthe inlet airflow causes the Mach number of the inlet airflow, to behigher than the Mach number of the freestream airflow. This differenceresults in an undesirable increase of a Thrust Specific Fuel Consumption(TSFC) of the unducted turbine engine when compared to an instance wherethe Mach number of the freestream airflow is equal to or less than theMach number of the inlet airflow. In other words, additional thrust andtherefore fuel consumption is required to accommodate for theaccelerated inlet airflow when compared to a scenario where the inletairflow is not accelerated relative to the freestream Mach number. It isdesirable to reduce the Mach number of the inlet airflow relative to thefreestream airflow to reduce the TSFC with respect to an acceleratedinlet airflow. It is desirable that the inlet airflow Mach number beequal to or less than the freestream Mach number.

One solution for controlling the Mach number of the inlet airflow (e.g.,ensure that the inlet airflow does not accelerate) is to mount theunducted turbine engine further from the fuselage by extending the pylonthat interconnects the unducted turbine engine and the fuselage of theaircraft. However, such extended pylons create additional structuralproblems related to the larger bending moments on the cantileveredpylon, which requires a more robust, and undesirably heavier, pylon. Theaddition of weight ultimately results in a heavier aircraft, whichnegatively affects the overall aircraft efficiency or performance,especially increasing the fuel consumption. The farther that theunducted turbine engine is extended from the fuselage centerline, thegreater will be the force effects, such as a yawing moment, associatedwith a thrust generated by the unducted turbine engine. In instanceswhere there are multiple unducted turbine engines on the aircraft, theseforce effects ultimately require larger control surfaces (more weight)to appropriately operate the aircraft in a single engine mode, if oneengine were to shut down.

Another solution for controlling the inlet airflow Mach number is toprovide a divot along the fuselage near the unducted turbine engine,without increasing the distance the unducted turbine engine is spacedfrom the fuselage centerline. Indeed, it was found that the size of adivot, the shape of the divot, and the positioning of the divot withrespect to the unducted turbine engine results in the desired control ofthe inlet airflow.

The process of designing a divot and the relative spacing of theunducted turbine engine and the divot, however, requires the formationof models or the use of simulations in order to determine whether or notthe divot and unducted turbine engine has the desired effect on theinlet airflow. This is an iterative, time consuming and costly processas the relative placement of the unducted turbine engine and the divot,along with the size of the divot, and selected based on an educatedguess, and then the Mach number of the inlet airflow calculated usingthe models and simulations, to assess the performance of the divot. Ifthe educated guess does not yield the desired performance result, theeducated guess is updated and the iteration is repeated.

The inventors' practice has proceeded in the foregoing manner ofdesigning an aircraft to include a divot in the fuselage of an aircraftwith an unducted turbine engine, designing at least one of the fuselage,the divot and the unducted turbine engine, followed by several redesignsof the fuselage, the divot and the unducted turbine engine to arrive ata desired control of the Mach number of the inlet airflow, thencalculating and checking the amount of fuel burn and thrust, andrepeating the process, etc. during the design of several differentaircrafts and unducted turbine engines.

More specifically, the inventors' practice has proceeded by makingeducated guesses on how various changes to defined geometricrelationships between the unducted turbine engine, the divot, and otherportions of the aircraft, such as a tail wing, would affect the Machnumber of the inlet airflow. To determine the inlet airflow Mach number,tests were conducted using models or simulations, such as ComputationalFluid Dynamics (CFD) models, and/or the building of physical prototypesand testing properties in a wind tunnel.

It is desired to identify the geometric relationships between theunducted turbine engine, the divot, and other portions of the aircraft,such as a tail wing, which yield advantageous TSFC ranges for aparticular aircraft construction and engine configuration. The desiredgeometric relationship would keep the inlet airflow Mach number within adesired range to achieve a desired TSFC operating range. In most cases,the desired range is where the inlet airflow Mach number is less thanthe freestream Mach number. It would also be desirable to arrive at thisadvantageous dimensional sizing of a divot and spacing without having togo through the time consuming and iterative process. That is, it isdesirable to establish conditions or limitations on the aircraftincluding the unducted turbine engine that account for the accelerationof the inlet airflow and the overall efficiency or robustness of theunducted turbine engine.

FIG. 1 is a schematic cross-sectional diagram of a turbine engine,specifically an unducted turbine engine 10 for an aircraft. The unductedturbine engine 10 has a generally longitudinally extending axis orengine centerline 12 extending from a forward end 14 to an aft end 16.The unducted turbine engine 10 includes, in downstream serial flowrelationship, a set of circumferentially spaced blades or propellersdefining a fan section 18 including a fan 20, a compressor section 22including a booster or low pressure (LP) compressor 24 and a highpressure (HP) compressor 26, a combustion section 28 including acombustor 30, a turbine section 32 including a HP turbine 34, and a LPturbine 36, and an exhaust section 38. The unducted turbine engine 10 asdescribed herein is meant as a non-limiting example, and otherarchitectures are possible, such as, but not limited to, the steamturbine engine, the supercritical carbon dioxide turbine engine, or anyother suitable turbine engine.

An exterior surface, defined by a nacelle 40, of the unducted turbineengine 10 extends from the forward end 14 of the unducted turbine engine10 toward the aft end 16 of the unducted turbine engine 10 and covers atleast a portion of the compressor section 22, the combustion section 28,the turbine section 32, and the exhaust section 38. The fan section 18can be positioned at a forward portion of the nacelle 40 and extendradially outward from the nacelle 40 of the unducted turbine engine 10,specifically, the fan section 18 extends radially outward from thenacelle 40. The fan section 18 includes a set of fan blades 42, and aset of stationary fan vanes 82 downstream the set of fan blades 42, bothdisposed radially about the engine centerline 12. The unducted turbineengine 10 includes any number of one or more sets of rotating blades orpropellers (e.g., the set of fan blades 42) disposed upstream of the setof stationary fan vanes 82. As a non-limiting example, the unductedturbine engine 10 can include multiple sets of fan blades 42 or fanvanes 82. As such, the unducted turbine engine 10 is further defined asan unducted single-fan turbine engine. The unducted turbine engine 10 isfurther defined by the location of the fan section 18 with respect tothe combustion section 28. The fan section 18 can be upstream,downstream, or in-line with the axial positioning of the combustionsection 28.

The compressor section 22, the combustion section 28, and the turbinesection 32 are collectively referred to as an engine core 44, whichgenerates combustion gases. The engine core 44 is surrounded by anengine casing 46, which is operatively coupled with a portion of thenacelle 40 of the unducted turbine engine 10.

A HP shaft or spool 48 disposed coaxially about the engine centerline 12of the unducted turbine engine 10 drivingly connects the HP turbine 34to the HP compressor 26. A LP shaft or spool 50, which is disposedcoaxially about the engine centerline 12 of the unducted turbine engine10 within the larger diameter annular HP spool 48, drivingly connectsthe LP turbine 36 to the LP compressor 24 and fan 20. The spools 48, 50are rotatable about the engine centerline 12 and couple to a set ofrotatable elements, which collectively define a rotor 51.

It will be appreciated that the unducted turbine engine 10 is either adirect drive or integral drive engine utilizing a reduction gearboxcoupling the LP shaft 50 to the fan 20. In some embodiments this gearboxhas a gear ratio of 4:1 or greater, for example, between 4:1 and 10:1and between 5:1 and 9:1. In some embodiments the engine 10 has a diskloading (defined as horsepower over swept area of the fan blades 42) ofbetween 60 and 150 HP/ft² and advance ratio above 3.8, as this term isdefined in US20210108572, herein incorporated by reference. The unductedturbine engine 116 preferable operates at a cruise speed Mach number ofgreater than or equal to 0.75 and less than or equal to 0.85, greaterthan or equal to 0.55, or greater than or equal to 0.75 and less than orequal to 0.85 depending on mission requirements.

The LP compressor 24 and the HP compressor 26, respectively, include aset of compressor stages 52, 54, in which a set of compressor blades 56,58 rotate relative to a corresponding set of static compressor vanes 60,62 (also called a nozzle) to compress or pressurize the stream of fluidpassing through the stage. In a single compressor stage 52, 54, multiplecompressor blades 56, 58 are provided in a ring and extend radiallyoutwardly relative to the engine centerline 12, from a blade platform toa blade tip, while the corresponding static compressor vanes 60, 62 arepositioned upstream of and adjacent to the compressor blades 56, 58. Itis noted that the number of blades, vanes, and compressor stages shownin FIG. 1 were selected for illustrative purposes only, and that othernumbers are possible.

The compressor blades 56, 58 for a stage of the compressor are mountedto a disk 61, which is mounted to the corresponding one of the HP and LPspools 48, 50, with each stage having its own disk 61. The staticcompressor vanes 60, 62 for a stage of the compressor are mounted to theengine casing 46 in a circumferential arrangement.

The HP turbine 34 and the LP turbine 36, respectively, include a set ofturbine stages 64, 66, in which a set of turbine blades 68, 70 arerotated relative to a corresponding set of static turbine vanes 72, 74(also called a nozzle) to extract energy from the stream of fluidpassing through the stage. In a single turbine stage 64, 66, multipleturbine blades 68, 70 are provided in a ring and extends radiallyoutwardly relative to the engine centerline 12, from a blade platform toa blade tip, while the corresponding static turbine vanes 72, 74 arepositioned upstream of and adjacent to the turbine blades 68, 70. It isnoted that the number of blades, vanes, and turbine stages shown in FIG.1 were selected for illustrative purposes only, and that other numbersare possible.

The turbine blades 68, 70 for a stage of the turbine are mounted to adisk 71, which is mounted to the corresponding one of the HP and LPspools 48, 50, with each stage having a dedicated disk 71. The staticturbine vanes 72, 74 for a stage of the compressor are be mounted to theengine casing 46 in a circumferential arrangement.

Complementary to the rotor portion, the stationary portions of theunducted turbine engine 10, such as the static vanes 60, 62, 72, 74among the compressor and turbine sections 22, 32 are also referred toindividually or collectively as a stator 63. As such, the stator 63refers to the combination of non-rotating elements throughout theunducted turbine engine 10.

The nacelle 40 is operatively coupled to the unducted turbine engine 10and covers at least a portion of the engine core 44, the engine casing46, or the exhaust section 38. At least a portion of the nacelle 40extends axially forward or upstream the illustrated position. Forexample, the nacelle 40 extends axially forward such that a portion ofthe nacelle 40 overlays or covers a portion of the fan section 18 or abooster section (not illustrated) of the unducted turbine engine 10.

During operation of the unducted turbine engine 10, a freestream airflow79 flows against a forward portion of the unducted turbine engine 10. Aportion of the freestream airflow 79 enters an annular area 25 definedby the swept area between the outer surface of the nacelle and the tipof the blade, with this air flow being an inlet airflow 78. A portion ofthe inlet airflow 78 enters the engine core 44 and is described as aworking airflow 76, which is used for combustion within the engine core44.

More specifically, the working airflow 76 flows into the LP compressor24, which then pressurizes the working airflow 76 thus defining apressurized airflow that is supplied to the HP compressor 26, whichfurther pressurizes the air. The working airflow 76, or the pressurizedairflow, from the HP compressor 26 is mixed with fuel in the combustor30 and ignited, thereby generating combustion gases. Some work isextracted from these gases by the HP turbine 34, which drives the HPcompressor 26. The combustion gases are discharged into the LP turbine36, which extracts additional work to drive the LP compressor 24, andthe working airflow 76, or exhaust gas, is ultimately discharged fromthe unducted turbine engine 10 via the exhaust section 38. The drivingof the LP turbine 36 drives the LP spool 50 to rotate the fan 20 and theLP compressor 24. The working airflow 76, including the pressurizedairflow and the combustion gases, defines a working airflow that flowsthrough the compressor section 22, the combustion section 28, and theturbine section 32 of the unducted turbine engine 10.

The inlet airflow 78 flows through the set of fan blades 42 and over thenacelle 40 of the unducted turbine engine 10. Subsequently, the inletairflow 78 flows over at least a portion of the set of stationary fanvanes 82, which directs the inlet airflow 78 such that it is transversetoward the engine centerline 12. The inlet airflow 78 then flows pastthe set of stationary fan vanes 82, following the curvature of thenacelle 40 and toward the exhaust section 38. A pylon 84 mounts theunducted turbine engine 10 to an exterior structure (e.g., a fuselage ofan aircraft, a wing, a tail wing, etc.).

The working airflow 76 and at least some of the inlet airflow 78 mergedownstream of the exhaust section 38 of the unducted turbine engine 10.The working airflow 76 and the inlet airflow 78, together, form anoverall thrust of the unducted turbine engine 10.

It is contemplated that a portion of the working airflow 76 is drawn asbleed air (e.g., from the compressor section 22). The bleed air providesan airflow to engine components requiring cooling. The temperature ofthe working airflow 76 exiting the combustor 30 is significantlyincreased with respect to the working airflow 76 within the compressorsection 22. As such, cooling provided by the bleed air is necessary foroperating of such engine components in the heightened temperatureenvironments or a hot portion of the unducted turbine engine 10. In thecontext of a turbine engine, the hot portions of the engine are normallydownstream of the combustor 30, especially the turbine section 32, withthe HP turbine 34 being the hottest portion as it is directly downstreamof the combustion section 28. Other sources of cooling fluid are, butare not limited to, fluid discharged from the LP compressor 24 or the HPcompressor 26.

FIG. 2 is a schematic perspective view of an aircraft 100 including ageneric unducted turbine engine 116 suitable for use as the unductedturbine engine 10 of FIG. 1 . The aircraft 100 includes a fuselage 102with an exterior surface. At least one wing 106 and a tail wing 108extend from the fuselage. The tail wing 108 is operably coupled to andspaced from the fuselage 102 via a tail wing pylon 114. The unductedturbine engine 116 is operably coupled to the exterior surface of thefuselage 102 via a pylon 120. The unducted turbine engine 116 includes aset of circumferentially spaced fan blades 118. A set of stationaryvanes 119 is provided downstream of the set of circumferentially spacedfan blades 118. The fuselage 102 extends between a nose 103 and a tail105 and includes a fuselage centerline 104 extending therebetween. Adivot 122 is formed along a portion of the exterior surface of thefuselage 102 near the unducted turbine engine 116. While illustrated asa single unducted turbine engine 116 and a single divot 122, it will beappreciated that there can be any number of unducted turbine engines 116coupled to a variety of suitable portions of the fuselage 102, the atleast one wing 106 or the tail wing 108. Further, there can be anynumber of one or more divots 122 provided about the exterior surface ofthe fuselage 102.

It will be further appreciated that the unducted turbine engine 116 isable to be operably coupled to a variety of aircraft 100 fuselage 102that have the divot 122. Additionally, while the tail wing 108 is aT-wing tail wing (e.g., the tail wing 108 as illustrated), otherconventional tail wings are contemplated such as, a cruciform tail wing,an H-tail, a triple tail, a V-tail, an inverted tail, a Y-tail, atwin-tail, a boom-mounted tail, or a ring tail, all of which arereferred to herein as the tail wing 108.

FIG. 3 is a top-down view of a portion of the aircraft 100 as viewedfrom a horizontal plane above the aircraft 100. The divot 122 extendsfrom an upstream edge 130 to a downstream edge 132. The divot 122 formsa depression in the otherwise normal shape of the fuselage 102. Thephysical relationship between the divot 122 and the unducted turbineengine 116 is quantified by using a set of geometric relationships.

As it moves axially, relative to the fuselage centerline, the divot 122converges radially inwardly from the upstream edge 130 to a maximumdepth 134, and then diverges radially outwardly from the maximum depth134 to the downstream edge 132. The rate of convergence to the maximumdepth 134 or rate of divergence from the maximum depth need not be thesame, and need not be constant. As can be seen, the divot 122 convergesto the maximum depth 134 at a first rate or slope, and diverges from themaximum depth 134 at a second rate or slope, which is different from thefirst rate or slope. The divot 122 is non-symmetric or symmetric about avertical plane normal to the horizontal plane as viewed in FIG. 2 , withthe vertical plane intersecting the maximum depth 134. The divot 122includes a variety of suitable shapes and sizes. The upstream edge 130and the downstream edge 132 are be provided along a variety of suitableportions of the fuselage 102.

The unducted turbine engine 116 is operably coupled to a suitablelocation along the fuselage 102 via the pylon 120. As illustrated, thepylon 120 is coupled to a portion of the fuselage 102 defined by thedivot 122. Alternatively, the pylon 120 is coupled to a portion of thefuselage 102 not defined by the divot 122. The pylon 120 is defined by apylon centerline 121 extending between a portion of the pylon 120coupled to the fuselage 102 and a portion of the pylon 120 coupled tothe unducted turbine engine 116.

The unducted turbine engine 116 is defined by an engine centerline 136.Each fan blade of the set of circumferentially spaced fan blades 118extends between a leading edge 124 and a trialing edge 126 in the axialdirection with respect to the engine centerline 136 and between a root128 and a tip 127 in the radial direction with respect to the enginecenterline 136. Each fan blade of the set of circumferentially spacedfan blades 118 meet with an exterior portion (e.g., a nacelle) of theunducted turbine engine 10 at the root 128. The engine centerline 136 isparallel or non-parallel to the fuselage centerline 104 when viewedalong a horizontal plane extending along the fuselage centerline 104.

A straight reference line 142, when viewed from the horizontal plane ofFIG. 3 , connects the upstream edge 130 and the downstream edge 132 ofthe divot 122. The straight reference line 142 has a length (L).

A given axial point along the divot 122 has a depth or height that isdefined by a line perpendicular to the straight reference line 142 andintersecting the divot 122. A maximum height (h) occurs at the maximumdepth 134.

The tip 127 of each fan blade of the set of circumferentially spaced fanblades 118 is provided a radial distance from the engine centerline 136.During rotation, the set of circumferentially spaced fan blades 118sweep an area to define a circle with a diameter (D).

The set of circumferentially spaced fan blades 118 include an axiallyforwardmost point 138 when viewed along a horizontal plane extendingalong the engine centerline 136 and intersecting the set ofcircumferentially spaced fan blades 118. A first distance (x1) isdefined as the distance, along a line parallel to the fuselagecenterline 104, between the axially forwardmost point 138 and theupstream edge 130 of the divot 122.

The tail wing 108 includes a leading edge 152 defining an axiallyforward edge of the tail wing 108 with respect to the fuselagecenterline 104. As illustrated, the downstream edge 132 of the divot 122is axially aft of, fore of, or coincide with the leading edge 152 of thetail wing 108 with respect to the fuselage centerline 104. A seconddistance (x2) is defined as the distance, along a line parallel to thefuselage centerline 104, between an axially forwardmost point 151 of theleading edge 152 of the tail wing 108 and the axially forwardmost point138 of the set of circumferentially spaced fan blades 118.

In one implementation, the unducted turbine engine 116 is axiallypositioned, with respect to the fuselage centerline 104, along thefuselage 102 to correspond with the location of the divot 122. Theunducted turbine engine 116 is positioned such that the axiallyforwardmost point 138 is axially aft of or coincides with the maximumdepth 134.

FIG. 4 is the same view as FIG. 3 and illustrates the airflows duringthe operation of the aircraft 100 and unducted turbine engine 116.During operation of the aircraft 100, a freestream airflow 79 flows overthe fuselage 102. A portion of the freestream airflow 79 flows over andinto the annulus defined between the annular area swept by the set ofblades from the nacelle to the tips to form the inlet airflow 78. Theinlet airflow 78 defines a first thrust airflow 146, while the workingairflow defines a second thrust airflow 148, with the first and secondthrust airflows 146, 148 defining an overall thrust of the unductedturbine engine 116.

The speed of the freestream airflow 79 is defined by a freestreamairflow Mach number (Mf) and the speed of the inlet airflow 78 isdefined by an inlet airflow Mach number (Mi).

FIG. 5 is a front view of an aircraft 200, suitable for use as theaircraft 100 of FIG. 2 , as viewed from a vertical plane. The aircraft200 is similar to the aircraft 100, therefore, like parts will beidentified with like numerals increased to the 200 series, with it beingunderstood that the description of the like parts of the aircraft 100applies to the aircraft 200 unless otherwise noted.

The aircraft 200 includes a fuselage 202 with a fuselage centerline 204.A pair of wings 206 extend from the fuselage 202. A tail wing pylon 214operably couples a tail wing 208 to the fuselage 202. At least one divot222 is formed along the fuselage 202. Two unducted turbine engines 216are spaced from the fuselage 202. Each unducted turbine engine 216 isoperably coupled to the fuselage 202 via a respective pylon 220. Thepylon 220 is defined by a pylon centerline 221. Each unducted turbineengine 216 is defined by an engine centerline 236. While notillustrated, each unducted turbine engine 216 includes a set ofcircumferentially arranged fan blades that define a swept area defininga circle 256 having a diameter (D). A boundary of the circle 256corresponds to at least one tip of the set of circumferentially arrangedfan blades.

The aircraft 200 includes two unducted turbine engines 216, one on eachside of the fuselage 202. The pylon 220 couples the unducted turbineengine 216 to a portion of the fuselage 202 corresponding to the divot222. The pylon 220 extends from the fuselage 202 at a roll angle (α)formed between the pylon centerline 221 and a horizontal plane. The rollangle (α) is greater than or equal to 0 degrees and less than or equalto 45 degrees (00≤α≤45°). It will be appreciated that the roll angle (α)can be a negative angle such that the roll angle is greater than orequal to −45 degrees and less than or equal to 0 degrees (−45°≤α≤0°).

A first radial distance (w1) is defined as the distance along the pyloncenterline 221 between the boundary of the circle 256 and the fuselage202 when viewed in the vertical plane of FIG. 5 . A second radialdistance (w2) is defined as the shortest radial distance between theexterior of the tail wing pylon and boundary of the circle 256 along ahorizontal line passing through engine centerline 236 when viewed in thevertical plane of FIG. 5 . A third radial distance (w3) is defined asthe shortest radial distance between the boundary of the circle 256 andthe exterior surface of the tail wing 208 along a radial line extendingfrom the engine centerline 236, when viewed in the vertical plane ofFIG. 5 .

With reference to FIGS. 2-5 , TABLES I-III disclose several embodimentsof the unducted turbine engine 116, 216, with different mountinglocations and different sizes for the divot 122, 222. The embodiments inTABLES I-III illustrate the values for the various geometriccharacteristics used to quantify the physical relationship between thedivot 122, 222 and the unducted turbine engine 116, 216, respectively,along with the corresponding range of inlet airflow Mach numbers (Mi)for the simulated freestream Mach number (Mf).

In Table I, the diameter (D) is held constant for embodiments 1-4 whilethe other physical characteristics (e.g., the first distance (x1) andthe length (L) and the height (h)) are varied, as well as the freestreamairflow Mach number (Mf) being varied. The corresponding inlet airflowMach number (Mi) range according to the simulation is shown. Themeasurements in the TABLES below are in meters (m).

TABLE I Embodiment D (m) x1 (m) L (m) h (m) Mf Mi 1 3 2.7 6 0.72 0.550.45-0.549 2 3 3.6 7.5 0.75 0.65 0.55-0.649 3 3 4.68 9 0.72 0.750.65-0.749 4 3 5.78 10.5 0.74 0.85 0.75-0.849

In Table II, the diameter D is held constant for embodiments 5-8, whilethe physical characteristic x2 is varied, along with the freestream Machnumber (Mf), which results in a corresponding inlet airflow Mach number(Mi) range.

TABLE II Embodiment D (m) x2 (m) Mf Mi 5 3.5 2.45 0.55 0.45-0.549 6 3.52.8 0.65 0.55-0.649 7 3.5 3.15 0.75 0.65-0.749 8 3.5 3.5 0.85 0.75-0.849

In Table III, the diameter D was held constant for embodiments 9-12while the physical characteristics w1, w2, and w3 were varied, alongwith the freestream Mach number (Mf), resulting in the correspondinginlet airflow Mach number (Mi) range.

TABLE III Embodiment D (m) w1 (m) w2 (m) w3 (m) Mf Mi  9 3.2 1.12 2.561.6 0.55 0.45-0.549 10 3.2 0.96 2.24 1.12 0.65 0.55-0.649 11 3.2 0.81.92 0.8 0.75 0.65-0.749 12 3.2 0.64 1.6 0.32 0.85 0.75-0.849

In all cases, the embodiments listed in TABLES I-III result in an inletairflow Mach number (Mi) range that is less than the freestream airflowMach number (Mf). The inlet airflow Mach number (Mi) is reduced bybetween 0.001 and 0.1 with respect to the freestream airflow Mach number(Mf). The reduction of the inlet airflow Mach number (Mi) with respectto the corresponding freestream airflow Mach number (Mf) ultimatelyresults in a decrease of the TSFC of the unducted turbine engine 116,216 when compared to an aircraft where the inlet airflow Mach number(Mi) is greater than or equal to freestream Mach number (Mf) (e.g.,Mi≤Mf). It has been found that the reduction of the inlet airflow Machnumber (Mi) with respect to the freestream airflow Mach number (Mf)results in an up to 5% decrease in the TSFC with respect to an aircraftwhere Mi≥Mf.

FIG. 6 is a schematic front view, as seen through a vertical plane, ofan aircraft 300 with an alternative tail section than that depicted inFIG. 2 . The aircraft 300 is similar to the aircraft 100, 200,therefore, like parts will be identified with like numerals increased tothe 300 series, with it being understood that the description of thelike parts of the aircraft 100, 200 applies to the aircraft 300 unlessotherwise noted.

The aircraft 300 includes a fuselage 302 extending between a nose (notillustrated) and a tail 305. The fuselage 302 is defined by a fuselagecenterline 304. The aircraft has wings 306. An unducted turbine engine316 defined by an engine centerline 336 is coupled to the fuselage 302via a pylon 320. The unducted turbine engine 316 includes a set ofcircumferentially spaced fan blades 318, which, during rotation, sweepan area defining a circle 356 with a diameter (D). The fuselage 302includes a divot 322.

The aircraft 300 includes the geometric characteristics described interms of the aircraft 100 (FIGS. 2-4 ). The aircraft 300 furtherincludes additional geometric characteristics that are used in tandemwith or in place of the geometric characteristics of the aircraft 100 todetermine an advantageous positioning and sizing of the divot 322relative to the unducted turbine engine 316.

The aircraft 300 includes a tail wing 308 in a V-formation (a V-tail)extending from respective portions of the fuselage 302. The divot 322and the unducted turbine engine 316 are each provided within theinterior of the V-tail tail wing 308 with respect to the fuselagecenterline 304.

A fourth radial distance (w4), which is useful for the V-formation tailwing 308, is defined as the shortest radial distance, with respect tothe engine centerline 336, between an exterior of one tail of the tailwing 308 and the engine centerline 336, when viewed from a verticalplane looking down the engine centerline as seen in FIG. 6 . Theunducted turbine engine 116 is located in the middle of the two tailsthat define the tail wing 308, thus, the fourth radial distance (w4) isthe same for each tail.

TABLE IV discloses several embodiments of mounting of the unductedturbine engine 316 relative to a V-tail tail wing 308, the location ofthe divot 322, and the size of the divot 322. The diameter D is keptconstant while w4 is varied, along with the simulated freestream Machnumber (Mf), which yields the corresponding inlet airflow Mach number(Mi) range.

TABLE IV Embodiment D (m) w4 (m) Mf Me 13 3.5 2.45 0.55 0.45-0.549 143.5 3.5 0.65 0.55-0.649 15 3.5 4.03 0.75 0.65-0.749 16 3.5 4.38 0.850.75-0.849

FIG. 7 is a schematic side view of the aircraft 300 of FIG. 6 , whenviewed from a vertical plane, which is parallel to the fuselagecenterline 304, and further illustrating additional geometriccharacteristics. The divot 322 extends axially between an upstream edge330 and a downstream edge 332 with respect to the fuselage centerline304. The unducted turbine engine 316 includes a set of circumferentiallyspaced fan blades 318 that extend between a root 328 and a tip 327 andbetween a leading edge 324 and a trailing edge 326. A set of stationaryvanes 319 are provided downstream of the set of circumferentially spacedfan blades 318. An axially forwardmost point 338 of the set ofcircumferentially spaced fan blades 318 defines a forwardmost portion ofthe circumferentially spaced fan blades 318 with respect to the fuselagecenterline 304.

A third distance (x3) is defined as the axial distance, with respect tothe engine centerline 336, between the axially forwardmost point 338 anda point 390 where the engine centerline 336 intersects a leading edge352 of the tail wing 308, when viewed from the vertical plane of FIG. 7. A fourth distance (x4) is defined as the distance, along a line thatis parallel to the engine centerline 336, between the axiallyforwardmost point 338 and a point 392 where a leading edge 354 of thepylon 320 intersects the unducted turbine engine 316.

TABLE V discloses several embodiments of mounting of the unductedturbine engine 316 in a V-tail tail wing 308, the location of the divot322, and the size of the divot 322. In TABLE V, the diameter D is heldconstant while x3 and x4 are varied, along with the simulated freestreamMach number (Mf) for embodiments 17-20, which results in thecorresponding inlet airflow Mach number (Mi) range.

TABLE V Embodiment D (m) x3 (m) x4 (m) Mf Mi 17 3.5 0.7 0.53 0.550.45-0.549 18 3.5 1.4 0.88 0.65 0.55-0.649 19 3.5 2.1 1.23 0.750.65-0.749 20 3.5 2.8 1.58 0.85 0.75-0.849

During their work on the previously described embodiments, the inventorsdiscovered relationships between the geometric characteristics of adivot and an unducted turbine engine configuration, which resulted inthe desired freestream airflow Mach number (Mf) to inlet airflow Machnumber (Mi) relationship yielding an improved TSFC without sacrificingoverall thrust of the engine. This relationship between the engine,mounting and fuselage shaping near the engine resulted in an improvedTSFC. The discovered relations are also beneficial to reduce reliance onthe trial and error approach described earlier, thereby reducing timeand costs.

The discovered relationships between certain geometric characteristicsare used to obtain the desired inlet airflow Mach number (Mi), overallthrust of the turbine engine, and TSFC. These discovered relationshipsare quantified as specific flow ratios, or simply “FR”.

These flow ratios are a first flow ratio (FR1), a second flow ratio(FR2) and a third flow ratio (FR3):

$\begin{matrix}{{{FR}1} = \frac{h}{L}} \\{{{FR}2} = \frac{L}{D}} \\{{{FR}3} = \frac{x1}{L}}\end{matrix}$

The first flow ratio (FR1) is greater than or equal to 0.01 and lessthan or equal to 0.15 (0.01≤FR1≤0.15). The first flow ratio (FR1) isused to determine the overall length (L) and height (h) of the divot.

The second flow ratio (FR2) is greater than or equal to 0.6 and lessthan or equal to 4 (0.6≤FR2≤4). The second flow ratio (FR2) is used todetermine the overall length (L) of the divot or the sizing of thecircumferentially spaced fan blades (e.g., the diameter (D)). Forexample, if the design of the aircraft requires the unducted turbineengine to have a set diameter (D), then the second flow ratio (FR2) isused to determine how long the divot needs to be to meet the requiredMach number of the inlet airflow. With the length (L) set, the overallheight (h) is picked through use of the first flow ratio (FR1).

The third flow ratio (FR3) is greater than or equal to 0.25 and lessthan or equal to 0.75 (0.25≤FR3≤0.75). The third flow ratio (FR3) isused to axially locate the unducted turbine engine with respect to theupstream edge of the divot. For example, if the sizing of the divot(e.g., the length (L)), the positioning of the axially forwardmost pointis determined by calculating the first distance (x1).

TABLE VI illustrates the first flow ratio (FR1), the second flow ratio(FR2) and the third flow ratio (FR3) for the embodiments in TABLE I.

TABLE VI Embodiment D (m) x1 (m) L (m) h (m) Mf Me FR1 FR2 FR3 1 3 2.7 60.72 0.55 0.45-0.549 0.12 2 0.45 2 3 3.6 7.5 0.75 0.65 0.55-0.649 0.12.5 0.48 3 3 4.68 9 0.72 0.75 0.65-0.749 0.08 3 0.52 4 3 5.78 10.5 0.740.85 0.75-0.849 0.07 3.5 0.55

The first flow ratio (FR1), the second flow ratio (FR2), and the thirdflow ratio (FR3) are used to position the unducted turbine engine withrespect to the divot to achieve the desired inlet airflow Mach number(Mi) for a specific freestream airflow Mach number (Mf). Further,through the ratios described, various geometric characteristics of thedivot and/or the unducted turbine engine are able to be calculated ifone or more geometric characteristics are known. For example, if thesize (e.g., the diameter (D)) of the unducted turbine engine is known,the second flow ratio (FR2) is able to be used to find the length (L) ofthe divot, which then allows for the calculation of the height (h) andthe first distance (x1) through use of the first flow ratio (FR1) andthe third flow ratio (FR3). This process is able to be done for eachfreestream airflow Mach number (Mf).

A fourth flow ratio (FR4) is defined as follows:

${{FR}4} = \frac{x2}{D}$

The fourth flow ratio (FR4) is greater than or equal to 0.5 and lessthan or equal to 2 (0.5≤FR4≤2). In summary, TABLE VI illustrates thefourth flow ratio (FR4) using the embodiments from TABLE II.

TABLE VI Embodiment D (m) x2 (m) Mf Me FR4 5 3.5 2.45 0.55 0.45-0.5490.7 6 3.5 2.8 0.65 0.55-0.649 0.8 7 3.5 3.15 0.75 0.65-0.749 0.9 8 3.53.5 0.85 0.75-0.849 1

The fourth flow ratio (FR4) is used to position the unducted turbineengine with respect to the tail wing to achieve the desired inletairflow Mach number (Mi) for a specific freestream airflow Mach number(Mf). Further, through the ratios described, various geometriccharacteristics of the divot and/or the unducted turbine engine are ableto be calculated if one or more geometric characteristics are known. Forexample, if the size (e.g., the diameter (D)) of the unducted turbineengine is known, the fourth flow ratio (FR4) is able to be used tocalculate the second distance (x2) for a given freestream airflow Machnumber (Mf).

A fifth flow ratio (FR5), a sixth flow ratio (FR6), and a seventh flowratio (FR7) are defined by the following, respectively:

$\begin{matrix}{{{FR}5} = \frac{w1}{D}} \\{{{FR}6} = \frac{w2}{D}} \\{{{FR}7} = \frac{w3}{D}}\end{matrix}$

The fifth flow ratio (FR5) is greater than or equal to 0.1 and less thanor equal to 1 (0.1≤FR5≤1). The sixth flow ratio (FR6) is greater than orequal to 0.1 and less than or equal to 1 (0.1≤FR6≤1). The seventh flowratio (FR7) is greater than or equal to 0.1 and less than or equal to 1(0.1≤FR7≤1).

TABLE VIII illustrates the fifth flow ratio (FR5), the sixth flow ratio(FR6) and the seventh flow ratio (FR7) using the embodiments from TABLEIII.

TABLE VIII Embodiment D (m) w1 (m) w2 (m) w3 (m) Mf Me FR5 FR6 FR7  93.2 1.12 2.56 1.6 0.55 0.45-0.549 0.35 0.8 0.5 10 3.2 0.96 2.24 1.120.65 0.55-0.649 0.3 0.7 0.35 11 3.2 0.8 1.92 0.8 0.75 0.65-0.749 0.250.6 0.25 12 3.2 0.64 1.6 0.32 0.85 0.75-0.849 0.2 0.5 0.1

The fifth flow ratio (FR5), the sixth flow ratio (FR6) and the seventhflow ratio (FR7) are used to position the unducted turbine engine withrespect to the tail wing and the divot to achieve the desired inletairflow Mach number (Mi) for a specific freestream airflow Mach number(Mf). Further, through the ratios described, various geometriccharacteristics of the divot and/or the unducted turbine engine are ableto be calculated if one or more geometric characteristics are known. Forexample, if the size (e.g., the diameter (D)) of the unducted turbineengine is known, the fifth flow ratio (FR5), the sixth flow ratio (FR6)and the seventh flow ratio (FR7) are able to be used to calculate thefirst radial distance (w1), the second radial distance (w2) and thethird radial distance (w3), respectively, for a given freestream airflowMach number (Mf).

An eighth flow ratio (FR8) is defined by the following:

${FR8} = \frac{w4}{D}$

The eighth flow ratio (FR8) is greater than or equal to 0.55 and lessthan or equal to 1.5 (0.55≤FR8≤1.5). In summary, TABLE IX illustratesthe eighth flow ratio (FR8) using the embodiments of TABLE IV.

TABLE IX Embodiment D (m) w4 (m) Mf Me FR8 13 3.5 2.45 0.55 0.45-0.5490.7 14 3.5 3.5 0.65 0.55-0.649 1 15 3.5 4.03 0.75 0.65-0.749 1.15 16 3.54.38 0.85 0.75-0.849 1.25

The eighth flow ratio (FR8) is used to position the unducted turbineengine with respect to the tail wing to achieve the desired inletairflow Mach number (Mi) for a specific freestream airflow Mach number(Mf). Further, through the ratios described, various geometriccharacteristics of the divot and/or the unducted turbine engine are ableto be calculated if one or more geometric characteristics are known. Forexample, if the size (e.g., the diameter (D)) of the unducted turbineengine is known, the eight flow ratio (FR8) is able to be used tocalculate the fourth radial distance (w4) for a given freestream airflowMach number (Mf).

A ninth flow ratio (FR9) and a tenth flow ratio (FR10) are defined bythe following, respectively:

$\begin{matrix}{{{FR}9} = \frac{x3}{D}} \\{{{FR}10} = \frac{x4}{D}}\end{matrix}$

The ninth flow ratio (FR9) is greater than or equal to 0.1 and less thanor equal to 1 (0.1≤FR9≤1). The tenth flow ratio (FR10) is greater thanor equal to 0.1 and less than or equal to 0.5 (0.1≤FR10≤0.5). Insummary, TABLE X illustrates the ninth flow ratio (FR9) and the tenthflow ratio (FR10) using the embodiments of TABLE V.

TABLE X D x3 x4 Embodiment (m) (m) (m) Mf Me FR9 FR10 17 3.5 0.7 0.530.55 0.45-0.549 0.2 0.15 18 3.5 1.4 0.88 0.65 0.55-0.649 0.4 0.25 19 3.52.1 1.23 0.75 0.65-0.749 0.6 0.35 20 3.5 2.8 1.58 0.85 0.75-0.849 0.80.45

The ninth flow ratio (FR9) and the tenth flow ratio (FR10) are used toposition the unducted turbine engine with respect to the tail wing andthe pylon to achieve the desired inlet airflow Mach number (Mi) for aspecific freestream airflow Mach number (Mf). Further, through theratios described, various geometric characteristics of the divot and/orthe unducted turbine engine are able to be calculated if one or moregeometric characteristics are known. For example, if the size (e.g., thediameter (D)) of the unducted turbine engine is known, the ninth flowratio (FR9) and the tenth flow ratio (FR10) are able to be used tocalculate the third distance (x3) and the fourth distance (x4),respectively, for a given freestream airflow Mach number (Mf).

The ranges of the first flow ratio (FR1), the second flow ratio (FR2),the third flow ratio (FR3), the fourth flow ratio (FR4), the fifth flowratio (FR5), the sixth flow ratio (FR6), the seventh flow ratio (FR7),the eight flow ratio (FR8), the ninth flow ratio (FR9), and the tenthflow ratio (FR10) are used together, or independently of one another, todefine the size and shape of the divot, and the size and placement ofthe unducted turbine engine with respect to various portions of thedivot, the unducted turbine engine and/or the tail wing. This, in turn,is used to reduce the inlet airflow Mach number (Mi) with respect to thefreestream airflow Mach number (Mf). This ultimately results in a lowerTSFC and a more efficient engine without sacrificing the overall thrustof the unducted turbine engine.

Benefits associated with the aircraft and the flow ratios describedherein include a greater ease in design when compared to the design of aconventional aircraft including an unducted turbine engine. For example,the conventional aircraft is designed through the time-consuming, costlyand iterative process described previously. The iterative process is notguaranteed to produce an aircraft that falls within the required TSFC orthrust of the unducted turbine engine. The flow ratios as describedherein, however, always results in an aircraft that falls within thedesired TSFC and overall thrust of the unducted turbine engine asmeasured by the reduction of the inlet airflow Mach number with respectto the freestream airflow Mach number.

This written description uses examples to describe aspects of thedisclosure described herein, including the best mode, and also to enableany person skilled in the art to practice aspects of the disclosure,including making and using any devices or systems and performing anyincorporated methods. The patentable scope of aspects of the disclosureis defined by the claims, and may include other examples that occur tothose skilled in the art. Such other examples are intended to be withinthe scope of the claims if they have structural elements that do notdiffer from the literal language of the claims, or if they includeequivalent structural elements with insubstantial differences from theliteral languages of the claims.

Further aspects of the disclosure are provided by the subject matter ofthe following clauses:

An aircraft comprising a fuselage defining a fuselage centerline, thefuselage comprising a divot having an upstream edge and a downstreamedge axially aft of the upstream edge with respect to the fuselagecenterline, wherein the divot is defined by a straight reference lineconnecting the upstream edge and the downstream edge, with the straightreference line extending a length (L), and the divot having a maximumdepth (h) relative to the straight reference line, and an unductedturbine engine operably coupled to the fuselage, the unducted turbineengine comprising an engine core defining an engine centerline, anacelle circumscribing at least a portion of the engine core, and a setof blades operably coupled to at least a portion of the engine core, theset of blades having an axially forwardmost point and wherein the set ofblades defining a swept area defining a circle having a diameter (D),wherein a first flow ratio (FR1) of the aircraft is equal to:h/L, andwherein the first flow ratio (FR1) is greater than or equal to 0.01 andless than or equal to 0.15 (0.01≤FR1≤0.15).

An aircraft comprising a fuselage defining a fuselage centerline, a tailwing coupled to the fuselage and having a leading edge, with an axiallyforwardmost point relative to the fuselage centerline, an unductedturbine engine operably coupled to the fuselage, the unducted turbineengine comprising an engine core defining an engine centerline, anacelle circumscribing at least a portion of the engine core, and a setof blades operably coupled to at least a portion of the engine core, theset of blades having an axially forwardmost point, wherein the set ofblades define a swept area defining a circle having a diameter (D),wherein a second distance (x2) is a distance along a line parallel tothe fuselage centerline between the axially forwardmost point of theleading edge of the tail wing and the axially forwardmost point of theset of blades, wherein a flow ratio (FR4) is equal to:

$\frac{x2}{D},$

and wherein the flow ratio (FR4) is greater than or equal to 0.5 andless than or equal to 2 (0.5≤FR4≤2).

The aircraft of any preceding clause, wherein a second flow ratio (FR2)is equal to:L/D, and the second flow ratio (FR2) is greater than orequal to 0.6 and less than or equal to 4 (0.6≤FR2≤4).

The aircraft of any preceding clause, wherein a first distance (x1) is adistance along a line parallel to the fuselage centerline and betweenthe axially forwardmost point of the set of blades and the upstream edgeof the divot, and a third flow ratio (FR3) is equal to:

$\frac{x1}{L},$

and the third flow ratio (FR3) is greater than or equal to 0.25 and lessthan or equal to 0.75 (0.25≤FR3≤0.75).

The aircraft of any preceding clause, further comprising a pylondefining a pylon centerline and coupling the fuselage and the unductedturbine engine, wherein the pylon centerline defines a roll angle (α)with a horizontal plane, and wherein the roll angle (α) is greater thanor equal to 0 degrees and less than or equal to 45 degrees (00≤α≤45°).

The aircraft of claim 4, wherein at least a portion of the pylonintersects the divot.

The aircraft of any preceding clause, wherein a first radial distance(w1) is a distance along the pylon centerline between an outer boundaryof the circle and the fuselage, a fifth flow ratio (FR5) is equal to:

$\frac{w1}{D},$

and the fifth flow ratio (FR5) is greater than or equal to 0.1 and lessthan or equal to 1 (0.1≤FR5≤1).

The aircraft of any preceding clause, wherein a fourth distance (x4) isthe distance between the axially forwardmost point and the intersectionpoint when viewed along a vertical plane extending along the enginecenterline, a tenth flow ratio (FR10) is equal to:

$\frac{x4}{D},$

and the tenth flow ratio (FR10) is greater than or equal to 0.1 and lessthan or equal to 0.5 (0.1≤FR10≤0.5).

The aircraft of any preceding clause, further comprising a tail winghaving a leading edge, which includes an axially forwardmost point withrespect to the fuselage centerline.

The aircraft of any preceding clause, wherein a second distance (x2) isa distance along a line parallel to the fuselage centerline between theaxially forwardmost point for the set of fan blades and the axiallyforwardmost point for the leading edge of the tail wing, a fourth flowratio (FR4) is equal to:

$\frac{x2}{D},$

and the fourth flow ratio (FR4) is greater than or equal to 0.5 and lessthan or equal to 2 (0.5≤FR4≤2).

The aircraft of any preceding clause, wherein the tail wing comprises atail wing pylon extending between the tail wing and the fuselage, andwherein a second radial distance (w2) is the shortest radial distancebetween an exterior of the tail wing pylon and an outer boundary of thecircle along a horizontal plane passing through the engine centerline,and a sixth flow ratio (FR6) is equal to:

$\frac{w2}{D},$

and wherein the sixth flow ratio (FR6) is greater than or equal to 0.1and less than or equal to 1 (0.1≤FR6≤1).

The aircraft of any preceding clause, wherein a third radial distance(w3) is the shortest radial distance between an exterior of the tailwing and the boundary of the circle along the horizontal plane, and aseventh flow ratio (FR7) is equal to:

$\frac{w3}{D},$

and the seventh flow ratio (FR7) is greater than or equal to 0.1 andless than or equal to 1 (0.1≤FR7≤1).

The aircraft of any preceding clause, wherein a fourth radial distance(w4) is the shortest radial distance between an exterior of the tailwing and the engine centerline, an eighth flow ratio (FR8) is equal to:

$\frac{w4}{D},$

and the eighth flow ratio (FR8) is greater than or equal to 0.55 andless than or equal to 1.5 (0.55≤FR8≤1.5).

The aircraft of any preceding clause, wherein a third distance (x3) isthe axial distance between the axially forwardmost point and a pointwhere the engine centerline intersects the leading edge of the tailwing, a ninth flow ratio (FR9) is equal to:

$\frac{x3}{D},$

and the ninth flow ratio (FR9) is greater than or equal to 0.1 and lessthan or equal to 1 (0.1≤FR9≤1).

The aircraft of any preceding clause, wherein the divot is non-symmetricabout a vertical plane perpendicular to the fuselage centerline andintersecting the maximum depth.

The aircraft of any preceding clause, wherein the engine centerline isparallel to the fuselage centerline.

The aircraft of any preceding clause, wherein the straight referenceline extends non-parallel to the fuselage centerline.

The aircraft of any preceding clause, wherein during operation of theaircraft and the unducted turbine engine an inlet airflow flows againstthe axially forwardmost point of the set of blades, and the inletairflow includes a Mach number that is between 0.001 and 0.1 smallerthan a Mach number of a freestream airflow along the fuselage upstreamof the unducted turbine engine.

The aircraft of any preceding clause, wherein during operation of theaircraft and the unducted turbine engine a freestream airflow flowsalong the fuselage upstream of the unducted turbine engine, and whereinthe unducted turbine engine operates at a cruise speed Mach number ofthe freestream airflow greater than or equal to 0.55 and less than orequal to 0.85.

The aircraft of any preceding clause, wherein during operation of theaircraft and the unducted turbine engine an inlet airflow flows againstthe axially forwardmost point of the set of blades, and the inletairflow includes a Mach number that is between 0.001 and 0.1 smallerthan a Mach number of a freestream airflow along the fuselage upstreamof the unducted turbine engine.

What is claimed is:
 1. An aircraft comprising: a fuselage defining afuselage centerline, the fuselage comprising: a divot having an upstreamedge and a downstream edge axially aft of the upstream edge with respectto the fuselage centerline; wherein the divot is defined by a straightreference line connecting the upstream edge and the downstream edge,with the straight reference line extending a length (L), and the divothaving a maximum depth (h) relative to the straight reference line; andan unducted turbine engine operably coupled to the fuselage, theunducted turbine engine comprising: an engine core defining an enginecenterline; a nacelle circumscribing at least a portion of the enginecore; and a set of blades operably coupled to at least a portion of theengine core, the set of blades having an axially forwardmost point andwherein the set of blades defining a swept area defining a circle havinga diameter (D); wherein a first flow ratio (FR1) of the aircraft isequal to:h/L, and wherein the first flow ratio (FR1) is greater than orequal to 0.01 and less than or equal to 0.15 (0.01≤FR1≤0.15).
 2. Theaircraft of claim 1, wherein a second flow ratio (FR2) is equal to:L/D,and the second flow ratio (FR2) is greater than or equal to 0.6 and lessthan or equal to 4 (0.6≤FR2≤4).
 3. The aircraft of claim 1, wherein afirst distance (x1) is a distance along a line parallel to the fuselagecenterline and between the axially forwardmost point of the set ofblades and the upstream edge of the divot, and a third flow ratio (FR3)is equal to: $\frac{x1}{L},$ and the third flow ratio (FR3) is greaterthan or equal to 0.25 and less than or equal to 0.75 (0.25≤FR3≤0.75). 4.The aircraft of claim 1, further comprising: a pylon defining a pyloncenterline and coupling the fuselage and the unducted turbine engine;wherein the pylon centerline defines a roll angle (α) with a horizontalplane; and wherein the roll angle (α) is greater than or equal to 0degrees and less than or equal to 45 degrees (00≤α≤45°).
 5. The aircraftof claim 4, wherein at least a portion of the pylon intersects thedivot.
 6. The aircraft of claim 4, wherein a first radial distance (w1)is a distance along the pylon centerline between an outer boundary ofthe circle and the fuselage, a fifth flow ratio (FR5) is equal to:$\frac{w1}{D},$ and the fifth flow ratio (FR5) is greater than or equalto 0.1 and less than or equal to 1 (0.1≤FR5≤1).
 7. The aircraft of claim1, wherein a fourth distance (x4) is the distance between the axiallyforwardmost point and the intersection point when viewed along avertical plane extending along the engine centerline, a tenth flow ratio(FR10) is equal to: $\frac{x4}{D},$ and the tenth flow ratio (FR10) isgreater than or equal to 0.1 and less than or equal to 0.5(0.1≤FR10≤0.5).
 8. The aircraft of claim 1, further comprising a tailwing having a leading edge, which includes an axially forwardmost pointwith respect to the fuselage centerline.
 9. The aircraft of claim 8,wherein a second distance (x2) is a distance along a line parallel tothe fuselage centerline between the axially forwardmost point for theset of fan blades and the axially forwardmost point for the leading edgeof the tail wing, a fourth flow ratio (FR4) is equal to: $\frac{x2}{D},$and the fourth flow ratio (FR4) is greater than or equal to 0.5 and lessthan or equal to 2 (0.5≤FR4≤2).
 10. The aircraft of claim 8, wherein thetail wing comprises a tail wing pylon extending between the tail wingand the fuselage, and wherein: a second radial distance (w2) is theshortest radial distance between an exterior of the tail wing pylon andan outer boundary of the circle along a horizontal plane passing throughthe engine centerline; and a sixth flow ratio (FR6) is equal to:$\frac{w2}{D},$ and wherein the sixth flow ratio (FR6) is greater thanor equal to 0.1 and less than or equal to 1 (0.1≤FR6≤1).
 11. Theaircraft of claim 10, wherein a third radial distance (w3) is theshortest radial distance between an exterior of the tail wing and theboundary of the circle along the horizontal plane; and a seventh flowratio (FR7) is equal to: $\frac{w3}{D},$ and the seventh flow ratio(FR7) is greater than or equal to 0.1 and less than or equal to 1(0.1≤FR7≤1).
 12. The aircraft of claim 8, wherein a fourth radialdistance (w4) is the shortest radial distance between an exterior of thetail wing and the engine centerline, an eighth flow ratio (FR8) is equalto: $\frac{w4}{D},$ and the eighth flow ratio (FR8) is greater than orequal to 0.55 and less than or equal to 1.5 (0.55≤FR8≤1.5).
 13. Theaircraft of claim 8, wherein a third distance (x3) is the axial distancebetween the axially forwardmost point and a point where the enginecenterline intersects the leading edge of the tail wing, a ninth flowratio (FR9) is equal to: $\frac{x3}{D},$ and the ninth flow ratio (FR9)is greater than or equal to 0.1 and less than or equal to 1 (0.1≤FR9≤1).14. The aircraft of claim 1, wherein the divot is non-symmetric about avertical plane perpendicular to the fuselage centerline and intersectingthe maximum depth.
 15. The aircraft of claim 1, wherein the enginecenterline is parallel to the fuselage centerline.
 16. The aircraft ofclaim 1, wherein the straight reference line extends non-parallel to thefuselage centerline.
 17. The aircraft of claim 1, wherein duringoperation of the aircraft and the unducted turbine engine an inletairflow flows against the axially forwardmost point of the set ofblades, and the inlet airflow includes a Mach number that is between0.001 and 0.1 smaller than a Mach number of a freestream airflow alongthe fuselage upstream of the unducted turbine engine.
 18. The aircraftof claim 1, wherein during operation of the aircraft and the unductedturbine engine a freestream airflow flows along the fuselage upstream ofthe unducted turbine engine, and wherein the unducted turbine engineoperates at a cruise speed Mach number of the freestream airflow greaterthan or equal to 0.55 and less than or equal to 0.85.
 19. An aircraftcomprising: a fuselage defining a fuselage centerline; a tail wingcoupled to the fuselage and having a leading edge, with an axiallyforwardmost point relative to the fuselage centerline; an unductedturbine engine operably coupled to the fuselage, the unducted turbineengine comprising: an engine core defining an engine centerline; anacelle circumscribing at least a portion of the engine core; and a setof blades operably coupled to at least a portion of the engine core, theset of blades having an axially forwardmost point; wherein the set ofblades define a swept area defining a circle having a diameter (D);wherein a second distance (x2) is a distance along a line parallel tothe fuselage centerline between the axially forwardmost point of theleading edge of the tail wing and the axially forwardmost point of theset of blades; wherein a flow ratio (FR4) is equal to: $\frac{x2}{D};$and wherein the flow ratio (FR4) is greater than or equal to 0.5 andless than or equal to 2 (0.5≤FR4≤2).
 20. The aircraft of claim 19,wherein during operation of the aircraft and the unducted turbine enginean inlet airflow flows against the axially forwardmost point of the setof blades, and the inlet airflow includes a Mach number that is between0.001 and 0.1 smaller than a Mach number of a freestream airflow alongthe fuselage upstream of the unducted turbine engine.